An aircraft pilot adjusts the position (orientation) of surfaces such as rudders and elevators so as to control the aircraft's flight. In traditional aircraft designs, cables and other mechanical components directly link the control stick to the surfaces, usually with the help of actuators. Modern aircraft designs, on the other hand, feature complex electronic components within their Flight Control Systems, thereby introducing new variables and data processing platforms (computers, electronic circuitry) in the control loop absent from purely mechanical Flight Control Systems. At the same time that many new and useful functionalities are brought on, a more complex Flight Control System presents itself with the possibility of containing several distinctive sources of faulty behavior—such as, for example, wrong oscillatory inputs coming from a malfunctioning Flight Control Computer and sent to the control surfaces. Faulty persistent surface (e.g. rudder and/or elevator) oscillations while an aircraft is in air, when not correctly detected, may lead to structural damage due to cumulative loads that can exceed the aircraft designed oscillatory envelope and, eventually, may result in catastrophic events, especially in those cases when limit loads are repeatedly reached. In other words, under such persistent oscillations, control surfaces can permanently deform, crack or even snap off the aircraft.
Current methodologies for aircraft oscillatory fault detection due to Flight Control System malfunctions are generally based solely on surface oscillations data, i.e. the fault detection process typically takes into account only amplitudes, moments and frequencies of surface (rudder/elevator) oscillations as measured by sensors used to detect such oscillations.
In this sense, for instance, a method and device for detecting an overstepping of design loads of the fin of an aircraft caused by Flight Control System malfunctions is described by U.S. Pat. No. 7,271,741. In this document, “design loads” mean the maximum loads that can be supported by the fin without permanent deformations. Whether design loads are being overstepped can be determined by means of an assessment of its bending and twisting moments, which are simultaneously and constantly monitored during the flight of the aircraft, their values being compared with a safety envelope.
Methods based solely on surface oscillations—such as the one just mentioned—could be effective for medium and high frequencies oscillations, i.e. those in the range, for example, from 1 Hz to 20 Hz (approximately, depending on the aircraft). However, for low frequencies, i.e. those in the range between, for example, 0.1 Hz and 1 Hz (approximately, depending on the aircraft), there can be problems, namely:                It becomes difficult to clearly separate oscillatory commands caused by a Flight Control System fault from non-faulty oscillatory commands caused by the pilot (pilots' range of operation is between 0.1 Hz and 1 Hz). This means that a real fault may not be detected by this approach.        Rigid body dynamics that occur in frequencies from 0.1 Hz to 1 Hz may prevent this approach's compliance with fatigue and limit load avoidance requirements. This means that severe fault conditions, such as those leading to a limit load-reaching scenario, may not be properly detected.        
U.S. Pat. No. 5,319,296 describes an oscillatory servo-valve fault monitor that aims at identifying faults attributable to servo-control system components. Similarly to the method of U.S. Pat. No. 7,271,741, this approach also takes into account only the control surface behavior, i.e. the monitoring process can be regarded as a purely local one, not considering the overall aircraft operational status, namely: it does not assess the real time surface load behavior and it does not occupy itself with the structural impacts deriving from a persistent faulty oscillation before it is detected (i.e. regarded as a fault). Moreover, it does not clearly address fatigue life consumption criteria nor solid guidelines are provided as far as how to deal with a limit load-reaching event is concerned. Finally, and more importantly, the method is also not clear as to how oscillatory commands caused by a Flight Control System fault are not to be mistaken for non-faulty oscillatory commands caused by the pilot, which means that a real fault may not be detected by this method.
It hence becomes necessary to devise a new, highly reliable methodology to detect low frequency persistent oscillatory signals generated by a Flight Control System malfunction that does not allow nuisance fault detection events and does not fail to detect real fault events.
The exemplary illustrative non-limiting technology described herein provides a Tail Load Monitoring System that detects faulty low frequency (e.g. those in the range from 0.1 to 1 Hz) oscillatory conditions caused by a Flight Control System malfunction while the aircraft is in air by means of a continuous assessment of the estimated tail load behavior and some data processing. Estimation and data processing activities are provided by a dedicated architecture, i.e. a Tail Load Monitor.